Differential throttling has been used in the past for expendable launch vehicles having a multiple engine configuration and for linear aerospike engines designed for use in horizontal take-off, reusable launchers. An example of the latter is the X-33/VentureStar. However, to our knowledge, differential throttling has not been applied to vertical launch, single engine aerospike configurations despite the potential advantage. It appears that an important reason for this concerns the problems associated with thrust vector control and the combustion instability issues associated with the symmetrical annular combustors required by the current state of the art.
Patented prior art of potential interest includes U.S. Pat. No. 6,213,431 to Janeke which discloses a sonic aerospike rocket engine having a first fuel injector which is located at the leading end and which directs a first fuel towards the reaction plane. A second fuel injector is located in between the leading end and trailing end and directs a second fuel towards the reaction plane. A bell engine may be used in conjunction with the aerospike engine in outer space for optimal engine efficiency.
U.S. Pat. No. 6,205,770 to Williams et al. discloses a rocket engine which comprises first and second rotary injectors for injecting respective fuel and oxidizer propellant components into a first combustion chamber. The effluent therefrom drives a turbine that rotates the rotary injectors. The rotary injectors are adapted so as to isolate the low pressure propellant supply from the relatively high pressures in the respective combustion chambers.
U.S. Pat. No. 6,220,016 to Defever et al. discloses a rocket engine which comprises first and second combustion chambers with respective combustion chamber liners bounding respective annular passages. The first combustion chamber discharges into the second and the respective annular passages are in fluid communication with one another.
U.S. Pat. No. 5,622,046 to Michaels et al. discloses a multiple impinging stream vortex injector assembly wherein a multiple impinging stream vortex injector combines two mixing schemes into a single injector. Both first stage mixing or turbulent vortex mixing is accomplished by impinging momentum balanced, tangentially injected propellant streams onto one another.
U.S. Pat. No. 4,936,091 to Schoenman discloses a method for operating a rocket engine by injecting fuel and oxidizer into an elongated combustion chamber in two flows, viz., a core flow where the fuel and oxidizer are intimately mixed and immediately combusted, and a peripheral curtain flow which surrounds the core flow. The curtain flow is in contact with the combustion chamber wall to cool it and limit the heat transfer from the wall to the injector to prevent vapor locks in the injector.